The invention relates to control and damping of nutation of space vehicles which are spinning, i.e. have a rotational speed about one of their geometric axes, during at least certain phases of their missions.
By way of example, reference may be made to a satellite which, during its mission, first travels in a transfer orbit and then is brought by an apogee boost motor into a final geosynchronous orbit. As a general rule, the unit consisting of the satellite and its apogee boost motor is spin stabilized about the thrust axis of the motor before igniting the latter to inject the satellite into the final, for example geostationary, orbit. Some satellites are also spin stabilized in their final orbit.
It has been known for a long time that when a body rotates freely about its center of gravity, the geometrical axis (of maximum or minimum moment of inertia) about which the spinning is the highest, describes a movement of nutation about the direction of the momentum of the body. Nutation may develop under the effect of internal energy dissipations on board the body (slopping of liquids, flexibility of the structure, friction on bearings, for example, in the case of a space vehicle); it grows or decreases with time depending on whether the ratio .lambda. of the moment of inertia about the spin axis and the moment of inertia about transverse axes is less or greater than 1.
The amplitude of the nutation angle of a space vehicle should be controlled and restricted within limits which depend on the vehicle and the phase of the mission. Whenever it is possible to construct the satellite so that its constant spinning axis is the axis of maximum moment of inertia, damping of the nutation movement occurs naturally and it will be sufficient in general to enhance damping by using a passive damper. But practical requirements, such as the configuration of the launcher may lead one to construct the space vehicle so that the nutation will grow to such an extent that in the absence of an active control device, the vehicle would finally assume a "flat" movement of nutation, i.e. would tumble about an axis perpendicular to its nominal spin axis.
Active nutation control systems are already known which reduce the nutation angle by applying corrective torques. Typically, torquers using mass ejection are used. In such a case, the thrusters of the attitude control device of the vehicle are used for nutation control. The torque applied to the vehicle by the thrusters results in modification of the direction of the angular momentum of the vehicle in relation to an inertial reference. If the spin axis of the vehicle is the axis along which the thrust of the apogee boost motor is applied, it is essential that its angular direction be correct for achieving the required orbital parameters.
Consequently, it is important to minimize changes in the direction of the momentum (which direction is that of the spin axis of the vehicle after nutation damping) during the phase of nutation control and reduction of the amplitude of the nutation movement to an acceptable value.
To solve the problem on the satellite METEOSAT, a nutation damper was constructed which controls nutation by successive corrective actions in the form of torques, each due to several immediately successive pulses supplied by the thrusters, in directions such that they all tend to reduce the nutation, but approximately compensate each other insofar as the modification of the direction of the angular momentum of the satellite is concerned.
That approach is acceptable in the case of space vehicles whose ratio between the rotational and transverse moments of inertia is less than about 0.7. For configurations for which this ratio is between 0.7 and 1, the angular change of the momentum direction is very sensitive to the actuation of the nutation control means; for example, for certain ratios of moments of inertia, the change in the direction of the angular momentum due to an actuation sequence of the nutation control means is the sum of the shifts due to each actuation, and may become prohibitive for the mission.